1. Field of the Invention
The present invention relates generally to high speed missiles, and more particularly to a missile propelled by a solid rocket fuel engine to hypersonic velocities in which the size of the inlet opening is varied through the use of an axially translatable inlet plug, sensors to determine velocity, altitude, and location of the missile, and an on-board computer for continuously repositioning the plug in the inlet opening in response to the sensed conditions. As the missile reaches a predetermined trajectory position, the inlet opening is closed to reduce thrust, to diminish the forward speed of the missile, and to cause the missile to traverse a downward trajectory toward a target, while at the same time effecting a substantial reduction in the radar cross-section of the missile.
2. Description of Related Art
Conventional ramjet engine operation entails aerodynamically compressing air in an intake duct, burning a mixture of the compressed air and fuel in a combustion chamber, and then ejecting the resulting hot gases through a propulsion nozzle where the expanding gases are discharged at a high velocity.
Detachable rocket engines have been used for many years to provide sufficient acceleration for missiles so that a high enough speed can be achieved to compress air in the intake duct to begin operation of ramjet. After the missile reaches a high enough velocity for the ramjet to be self-sustaining, the rocket engine is detached and discarded.
To save space and reduce weight, rocket grain fuel has been provided within the ramjet combustion chamber for initial acceleration of the missile (see U.S. Pat. No. 3,452,544 to Glick et al., U.S. Pat. No. 3,609,977 to McCormick, and U.S. Pat. No. 4,891,938 to Nagy et al.).
Nevertheless, the integral rocket and ramjet engine design has drawbacks. For one thing, the engine still exhibits great weight. For another thing, the propelling nozzle inlet size must be changed when the engine converts from a rocket to a ramjet.
Addressing the weight issue has largely been a matter of reducing the number of components used or finding new or alternative materials for the various required engine components. Also, new fuel materials have been developed which either are of lower specific gravity or which require fewer support systems.
The problem of altering the nozzle inlet size has largely been approached by using different engine configurations, including ones in which the inlet opening is of a fixed size and a plug is axially moved into and out of blocking relationship with the opening (see U.S. Pat. No. 4,628,688 to Keirsey). The plug movement controls the consumption of fuel, as well as the transition from rocket engine operation to ramjet engine operation.
U.S. Pat. No. 2,684,570 to Nordfors shows one version of a missile which can transition from rocket engine operation to ramjet operation. The nose 4 initially blocks the inlet nozzle, permitting normal consumption of rocket fuel. When the fuel has been consumed, the reduced pressure behind the nose combined with pressure from oncoming ambient air forces the nose forced rearwardly, thereby "ramming" air into the ramjet engine.
Another technique for addressing this problem has been to burn away the rocket nozzle during transition (see U.S. Pat. No. 4,651,523 to Adams).
The excessive weight problem has been approached through the use of fewer components, or components made of lighter materials. This entails not only components but also body parts, such as support elements and skin. However, a significant difficulty in choosing a lighter material is that it must be able to stand up to extremely high temperatures whether from ramjet engine operation or from ambient friction experienced by the craft itself while traveling at hypersonic velocities.
For example, U.S. Pat. No. 5,594,216 to Yasukama et al. teaches using a light weight acoustic material comprised of a matrix of fused silica fibers as an insert in a jet engine housing.
U.S. Pat. No. 5,413,859 to Black et al. teaches a thermal protection system (TPS) for the nose tip of a reentry space vehicle which includes a carbon-carbon nose tip having a first and second sublimatable layers at inner and outer portions of the nose tip.
U.S. Pat. No. 5,560,569 to Schmidt discloses a thermal protection system for hypersonic cruise and space launch vehicles in which a flexible outer skin is formed from a metal super alloy secured over ceramic blocks that provide both an insulation layer and support for the outer skin. The blocks are made of a composite fiber-ceramic insulation material formed of silica and alumina fibers, boron nitride, and silicon carbide.
Notwithstanding that these disclosures might ultimately lead to an appropriate solution to the foregoing drawbacks, still another problem that has not been addressed by the known prior art is that of rendering an incoming missile as invisible as possible to radar and other missile detecting systems.
Against this background of known technology, the inventor has developed a missile propelled by solid rocket fuel to hypersonic velocities in which the size of the inlet opening is varied through the use of an axially translatable plug, and a computer carried on-board the missile determines the position of the plug in the inlet opening. In this manner, as the missile reaches a predetermined trajectory location and requires redirection toward a target, the inlet opening will be closed to reduce forward speed of the missile and facilitate its downward trajectory, while at the same time substantially minimizing the discernible radar cross-section of the missile.